1. Field of the Invention
This invention relates generally to gas turbine engines and, more specifically, to cooled turbine shrouds and impingement baffles that cool the shrouds.
2. Description of Related Art
In a gas turbine engine, air is pressurized in a compressor, mixed with fuel in a combustor, and ignited for generating hot combustion gases which flow downstream through one or more turbine stages which extract energy therefrom. A high pressure turbine (HPT) first receives the combustion gases from the combustor and extracts energy therefrom for powering the compressor. A low pressure turbine (LPT) downstream of the HPT extracts additional energy for providing output energy used for powering a fan disposed upstream of the compressor in a typical aircraft gas turbine engine application. In an industrial or a marine gas turbine engine, the LPT drives an output shaft for powering a generator or propellers of a ship. Shafts may also be used to drive helicopter blades or propellers of prop jet engines.
The HPT includes a stationary turbine nozzle having a plurality of circumferentially spaced apart stator vanes or turbine nozzles which control discharge of combustion gases from the combustor. The HPT also includes at least one rotor stage having a plurality of circumferentially spaced apart turbine rotor blades extending radially outwardly from a rotor disk. The blades include airfoils which receive combustion gases from the nozzle and extract energy therefrom for rotating the rotor disk and, in turn, rotating the compressor. The airfoils are typically hollow and include internal cooling circuits therein through which a portion of pressurized air bled from the compressor is channeled for cooling the blades.
Surrounding the rotor blades is an annular turbine shroud fixedly joined to a surrounding stator casing. The shroud is suspended closely atop the blade tips for providing a small gap or tip clearance therebetween. The tip clearance should be as small as possible to provide an effective fluid seal thereat during operation for minimizing the amount of combustion gas leakage therethrough for maximizing efficiency of operation of the engine. However, due to differential thermal expansion and contraction of the rotor blades and surrounding turbine shroud, the blade tips occasionally rub against the inner surface of the shroud causing abrasion wear.
Since the blade tips are at the radially outermost end of the rotor blades and are directly exposed to the hot combustion gases, they are difficult to cool and the life of the blade is thereby limited. The blade tips are typically in the form of squealer rib extensions of the pressure and suction sides of the airfoil, extending outwardly from a tip floor. Cooling air is channeled under the floor to cool the ribs by conduction and film cooling holes may extend through the floor to film cool the exposed ribs.
Since the turbine shroud is also exposed to the hot combustion gases, it too is also cooled by bleeding a portion of the pressurized air from the compressor, which is typically channeled in impingement cooling against the radially outer surface of the turbine shroud. Turbine shrouds typically also include film cooling holes extending radially therethrough with outlets on the radially inner surface of the shroud from which is discharged the cooling air in a film for cooling the inner surface of the shroud.
Impingement cooling of the shroud is also used. Baffles incorporate impingement cooling holes or apertures to direct cooling air against the back or radially outer surface of the shroud to achieve impingement cooling thereof. A relatively large amount of impingement cooling air is generally required for effective impingement cooling which decreases engine efficiency. Cooling air uses power from the engine and therefore causes the engine to use more fuel. Impingement cooling air is generally supplied to a plenum radially adjacent the shroud. The cooling air is supplied through inlet ports. The impingement holes are typically arranged in a circumferentially symmetric pattern with respect to an axis of rotation of arcuate shroud segments and corresponding baffles, thus, providing a substantially uniform circumferential discharge of the cooling air through the shroud.
In a high pressure turbine, the temperatures are not always circumferentially uniform and static components, such as turbine shrouds, in the flowpath can experience hot streaks. These hot streaks are due to the placement of combustor burners and also due to their location relative to static turbine nozzle airfoils. Pressure wakes from upstream turbine nozzle airfoils can locally reduce film and convective cooling in wake regions of the static component by reducing local pressure gradients and, thus, reduce film cooling air flow. The pressure wakes may also cause leakage flow of cooling air between a front face of the shroud and the upstream nozzle to be reduced in the high pressure areas, further increasing the local temperature of the shroud leading edge in this region. This effect may be particularly severe with new high performance nozzle designs incorporating 3-D aerodynamics and which are characterized by more severe pressure gradients at the flowpath edges. Local hot regions can result in oxidation and eventual burn through of the part. This can result in premature failure of the part and/or high scrap rates at overhaul.
It is desirable to provide a turbine shroud cooling system that accommodates circumferential heating gradients while minimizing the amount of cooling airflow, loss of engine efficiency, and fuel consumption.